JP3382786B2 - Rocket engine - Google Patents

Description

Description: BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a radiation-cooled rocket engine used for changing the orbit of a flying object such as an artificial satellite or controlling the attitude of the flying object. 2. Description of the Related Art FIG. 5 is a cross-sectional view of a rocket engine using a propellant using nitrogen tetroxide (NTO) as an oxidizing agent and monomethylhydrazine (MMH) as a fuel. Self-combustible NTO, MM that ignites only by contact of two liquids
In a rocket engine using H as a propellant, as shown in the figure, the fuel flows into the injector 02 from the axial direction of the combustion chamber 01, and flows into the injector 01 from the injected NTO and the radial direction of the combustion chamber 01, The injected MMH comes into direct contact and ignites, burns in the combustion chamber 01 at a combustion temperature of 3300 ° K or more, and this high-temperature, high-pressure combustion gas passes through the throat 03 and flows out of the nozzle 04 at high speed. The thrust is generated by ejecting the gas. In order to protect the nozzle 04, the throat 03, and the combustion chamber 01 (hereinafter, also referred to as a combustion chamber and the like) from a high combustion temperature, a passage 07 is provided inside these members.
MMH is injected from the inlet pipe 05, passed through the passage 07, cooled, and heated MMH1 flowing out of the outlet pipe 06 is discharged from the injector inlet 08 into the injector 0.
2 is adopted. [0004] However, an attitude control thruster engine (Reaction Control System) used for changing the orbit of an artificial satellite or the like or for attitude control.
In a rocket engine for RCS), the thrust is as small as about 200 N, and as shown in FIG. 6, a combustion chamber including a combustion chamber 01 ′, a throat 03 ′, and a nozzle 04 ′ are integrally formed. Is formed integrally with a valve that controls the injection amount of propellant, and the injector and the combustion chamber 01 'are integrated by welding without being joined by a flange,
The weight has been reduced. [0005] For this reason, the above-mentioned regenerative cooling system cannot be used for cooling a combustion chamber or the like, and a radiant cooling system is used. Further, in order to withstand the above-described high combustion temperature, the integrally formed combustion chamber and the like have ductility to withstand cold working, can be easily processed into a pipe or the like, and have a high melting point, high temperature strength and Niobium (N
iobium: Nb) is used. However, Nb is liable to be oxidized and forms an oxide film having a low melting point temperature, and is easily peeled off in a high-temperature atmosphere, resulting in a problem that the thickness is reduced. For this reason, as shown in FIG. 6, zirconium dioxide (ZrO ₂ ) is thermally sprayed 09 on the outer peripheral surface of a combustion chamber or the like made of Nb to perform oxidation-resistant coating or silicon carbide. SiC is deposited on a chemical vapor deposition solution (Chemical Vaccum Development).
ent: CVD) method, oxidation resistant coating,
That is, formation of a skin layer to prevent formation of an oxide film is performed. (1) However, in the method of spraying ZrO ₂ on the outer peripheral surface of such a combustion chamber made of Nb to form an oxidation resistant coating, these combustion chambers are used. Although surface temperatures up to 2000 ° C. have the advantage of being sufficiently usable, they have the disadvantage that they are easily peeled off and lack reliability because the oxidation resistant coating is formed by thermal spraying. (2) In the method of depositing SiC on the outer peripheral surface of the combustion chamber or the like made of Nb by the CVD method and coating it with oxidation resistance, there is an advantage that the coating is not peeled because of the coating by the CVD method. However, it has insufficient heat resistance, and can be used only at a surface temperature of 1400 ° C. to 1600 ° C., and cannot be used in a radiation cooling type combustion chamber or the like. SUMMARY OF THE INVENTION The present invention solves the above-mentioned problems of the rocket engine for RCS employing the conventional radiation cooling system. In addition, Ir, which has a high melting point and excellent oxidation resistance like Nb, is employed, and in order to compensate for the difference in the coefficient of thermal expansion between Nb and Ir, the concentration of the Nb and Ir layers is inclined. An object of the present invention is to provide a rocket engine in which the skin layer is hardly peeled off and has high durability by laminating the layers, and has improved heat resistance. [0010] Therefore, the rocket engine of the present invention has the following means. At least the structural members of the combustion chamber constituting the rocket engine are laminated as follows. (1) Niobium (Nb) having a high melting point of about 2750 ° K, excellent high-temperature strength, and a relatively small specific gravity of 8.57 is coated a plurality of times by a chemical (vacuum) vapor deposition (CVD) method. And laminated to strengthen the high-temperature strength. The strength layer is preferably formed by increasing the number of coatings by the chemical vapor deposition method, and desirably, to a thickness of about 1000 μm. (2) Iridium (Ir) having a high melting point of about 2730 ° K, excellent oxidation resistance, and a relatively large specific gravity of 22.5 is provided on the inner and outer peripheral surfaces of the strength layer. Nb is alternately laminated by coating by a CVD method, and at least two layers of Ir and Nb are provided on the inner peripheral surface and the outer peripheral surface of the strength layer.
And an intermediate layer that compensates for the difference in the coefficient of thermal expansion between the strength layer formed of and the skin layer formed of Ir. The layers of Ir and Nb laminated in the intermediate layer have a thickness of the order of several μm, and the lamination near the strength layer has a thick Nb layer, a thin Ir layer, and a thin skin layer. The lamination of the near part is I
It is preferable to form the layer so that the layer of r is thick and the layer of Nb is thin. (3) A skin layer for coating the inner and outer peripheral surfaces of an intermediate layer in which Ir and Nb layers are alternately laminated, by applying Ir a plurality of times by a CVD method, laminating the layers, and reinforcing the oxidation resistance. In addition, since the outer skin layer is formed of Ir having a higher specific gravity than Nb, it is preferable that the outer skin layer has a thickness of about 50 μm. According to the rocket engine of the present invention, at least the structural strength material of the combustion chamber, which becomes the highest temperature among the members exposed to the high-temperature combustion gas in the radiation-cooled rocket engine by the above means, Since it is formed by adopting a strength layer in which Nb is laminated, it is lightweight and has a heat resistance capable of sufficiently maintaining high-temperature strength even in a rocket engine having a surface temperature of 2000 ° C. using NTO or MMH as a propellant. It is rich in sex. Further, since the oxidation-resistant coating formed on the surface of the combustion chamber is formed by a skin layer laminated by using Ir, it can be used in a rocket engine having a surface temperature of 2000 ° C. In addition to being able to have high heat resistance, an oxide film is formed, the thickness of the oxide layer is reduced by peeling of the oxide film, and a decrease in the structural strength of the strength layer, which may be reduced in strength, can be prevented, and the durability can be enhanced. Further, the thermal expansion coefficient is 7.31 × 10 ⁻⁶ 1 /
The bonding between the strength layer made of Nb at ° K and the skin layer made of Ir at 6.8 × 10 ⁻⁶ 1 / ° K is made by an intermediate layer laminated by changing the distribution of Nb-Ir stepwise. So N
The difference in the coefficient of thermal expansion between b and Ir is reduced, so that no excessive stress is generated at the joint between the strength layer and the skin layer, the adhesion is improved, the peeling is difficult, and the durability is enhanced. it can. Incidentally, the above-mentioned means is adopted also for a throat and a nozzle which are usually formed integrally with the combustion chamber, and the same operation and effect as those described above can be obtained. An embodiment of a rocket engine according to the present invention will be described below with reference to the drawings. FIG. 1 is a longitudinal sectional view of a rocket engine for a thruster engine (RCS) for attitude control, which has a small thrust of about 200 N and is used for changing the orbit of an artificial satellite or the like or for attitude control according to a first embodiment of the present invention. FIG. As shown in the figure, the combustion chamber 1, the throat 3, and the nozzle 4 are integrated to form a combustion chamber or the like 5, and the tip of the combustion chamber 1 and the valve-integrated injector 2 are
It is directly welded without providing a flange or the like, thereby reducing the weight. Further, NTO as an oxidizing agent and MMH as a fuel are both introduced through tubes arranged substantially in parallel to the axial direction of the combustion chamber 1, and are introduced into a valve-integrated injector 2.
And is injected so as to collide with each other at the axis of the combustion chamber 1. Ignition by this collision, NTO, MMH
Is burned at a combustion temperature of 3000 ° C. or more in the combustion chamber 1, and this high-temperature, high-pressure combustion gas is
Through the throat 3, the thrust is generated by being accelerated by the nozzle 4 and ejected at high speed to the outside. On the other hand, FIG. 2 is a partial schematic cross-sectional view of a combustion chamber 1 and the like 5 composed of a combustion chamber 1, a throat 3, and a nozzle 4 formed in one body. As shown in FIG. 3, the combustion chamber 5 and the like are formed of the following five layers. (1) Niobium having a high melting point of about 2750 ° K, excellent high-temperature strength, a relatively small specific gravity of 8.57, and a coefficient of thermal expansion of 7.3 × 10 ^-6 1 / ° K. Nb)
Is coated a plurality of times by a chemical (vacuum) vapor deposition (CVD) method, laminated to a thickness of 1000 μm or more, and reinforced in strength. (2) Each of the inner and outer peripheral surfaces of the strength layer 6 has a high melting point of about 2730 ° K, excellent oxidation resistance, a relatively large specific gravity of 22.5, and a coefficient of thermal expansion. Is 6.8 × 10 ^-6 1 / ° K.
And Nb described above are alternately coated by a CVD method to be laminated to a thickness of the order of several μm, and a layer of Ir and Nb is formed on the inner peripheral surface and the outer peripheral surface of the strength layer 6,
At least two or more layers, and a strength layer 6 made of Nb
And an intermediate layer 7 which compensates for a difference in the coefficient of thermal expansion between a skin layer 8 formed of Ir and described later. (3) On the inner and outer peripheral surfaces of the intermediate layer 7,
A skin layer 8 that is coated with Ir a plurality of times by a CVD method and laminated to a thickness of about 50 μm to form an oxidation-resistant film, and prevents the formation of an oxide film on the strength layer 6 made of Nb. As shown in FIG. 4 which is a detailed view of the intermediate layer 7, the layers of Ir and Nb forming the intermediate layer 7 have a thickness near the strength layer 6 where the Nb layer is thickened. I
r is thinned, and the portion near the skin layer 8 is laminated as I
The r layer is formed to be thick and the Nb layer is formed to be thin. Next, a method of forming such a combustion chamber 5 composed of five layers will be described. First, aluminum A
The l-type mandrel 9 is processed in accordance with the inner surface shape of the combustion chamber 5 or the like, and its outer peripheral surface is coated with Ir, which is the material of the skin layer 8 forming the inner peripheral surface of the combustion chamber 5 by about 50 μm, by the CVD method. I do. In this case, since the specific gravity of Ir is as large as 22.5, it is necessary to prevent an increase in the weight of the combustion chamber 5 by taking care not to increase the thickness more than necessary. Next, Nb and Ir are deposited by CVD of about several μm.
The intermediate layer 7 is formed alternately by a coating method. At this time, as the number of layers increases, as shown in FIG.
Is made thicker and the Ir layer is made smaller. Then, on the outer peripheral surface of the minimum Ir layer of the intermediate layer 7, one layer of only Nb is formed.
000μm or more coated by CVD method, strength layer 6
To form Next, Ir and Nb are alternately coated by a CVD method to form an intermediate layer 7. The intermediate layer 7 formed on the outer peripheral surface side of the strength layer 6 has the structure shown in FIG.
The Nb layer is gradually made smaller and the Ir layer is made larger, contrary to the method shown in FIG. Above the minimum Nb layer of the intermediate layer 7, a layer of only Ir is coated by a few tens of μm by CVD to form the outermost skin layer 8. Finally, if the mandrel 9 made of Al is removed with caustic soda, the inner and outer surfaces become the combustion chamber 5 and the like on which the outer surface is made of Ir and the oxidation resistant coating is formed. As described above, in the rocket engine of the present embodiment, Ir is applied to Nb by oxidation-resistant coating by the CVD method, and further, in order to reduce the difference in thermal expansion coefficient between Ir and Nb,
Since the bonding was performed while changing the distribution of -Nb stepwise, Ir
The difference in thermal expansion between Nb and Nb is reduced, the junction between Ir and Nb is less likely to come off, and the durability is improved. Further, since the melting point of Ir is almost the same as that of Nb, the surface temperature is about 20 μm.
It can be used up to about 00 ° C. and has excellent heat resistance. The combustion chamber 1, the throat 3, and the nozzle 4 are integrated, and the regenerative cooling method cannot be used.
It can be applied to radiant cooling rocket engines that can reach 0 ° C. As described above, according to the rocket engine of the present invention, according to the configuration shown in the claims, (1) For example, NTO or MMH is used as a propellant,
Even in a rocket engine employing a radiation cooling system in which the surface temperature reaches 2000 ° C., the rocket engine can be made lightweight, have sufficient structural strength, and have high heat resistance. (2) A heat-resistant coating which can be used in a radiant cooling type rocket engine having a surface temperature as high as 2000 ° C. can be formed, and an oxide film is formed on the strength layer, and the thin oxide film is peeled off. And
A decrease in the structural strength of the strength layer, which may cause a decrease in strength, can be prevented, and the durability can be enhanced. (3) Further, Nb having a difference in thermal expansion coefficient,
The difference in the thermal expansion coefficient between Nb and Ir is reduced by the intermediate layer formed by changing the distribution of Nb-Ir in a stepwise manner in the bonding (stacking) of Ir, thereby improving the adhesion and making the bonded portion hard to come off. And can be made durable.

BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a longitudinal sectional view showing a first embodiment of a rocket engine of the present invention, FIG. 2 is a partial schematic sectional view of a combustion chamber, a throat and a nozzle of the embodiment shown in FIG. FIG. 3 is a cross-sectional view taken along the line AA shown in FIG. 2. FIG. 4 is a detailed cross-sectional view of the intermediate layer shown in FIG. 3. FIG. 5 is a rocket engine of a regenerative cooling system. FIG. 2 is a partial vertical cross-sectional view of the radiant cooling type rocket engine of FIG. [Description of Signs] 1 Combustion chamber 2 Injector 3 Throat 4 Nozzle 5 Combustion chamber etc. 6 Strength layer 7 Middle layer 8 Skin layer 9 Mandrel

────────────────────────────────────────────────── ─── Continuation of the front page (56) References JP-A-4-149082 (JP, A) JP-A-6-9269 (JP, A) JP-A-6-256926 (JP, A) (58) Field (Int.Cl. ⁷ , DB name) F02K 9/62 B64G 1/26 F02K 9/88

Claims (1)

(57) [Claims 1] In a radiation-cooled rocket engine used for changing the trajectory or controlling the attitude of a flying object, niobium is laminated a plurality of times by chemical vapor deposition to form a cylindrical shape. An intermediate layer in which at least two layers of the iridium and the niobium are formed by alternately laminating iridium and the niobium on the inner peripheral surface and the outer peripheral surface of the strength layer by chemical vapor deposition. A rocket engine, comprising: a combustion chamber including a skin layer formed by laminating the iridium a plurality of times by a chemical vapor deposition method on inner and outer peripheral surfaces of the intermediate layer.